U.S. patent number 8,317,128 [Application Number 12/932,091] was granted by the patent office on 2012-11-27 for laminar flow wing optimized for transonic cruise aircraft.
This patent grant is currently assigned to Aerion Corporation. Invention is credited to James D. Chase, Richard R. Tracy.
United States Patent |
8,317,128 |
Tracy , et al. |
November 27, 2012 |
Laminar flow wing optimized for transonic cruise aircraft
Abstract
On an aircraft designed for maximum efficient cruise speed in
the range from about Mach 0.8 to about Mach 1.2, and having
fuselage and wings with: (a) less than about 25 degrees of leading
edge sweep, in combination with airfoil thickness to chord ratios
between about 3% and about 8%, as an average along the wing
semi-span outboard from the zone of substantial fuselage influence,
and (b) wing leading edge sweep between about 20 degrees and about
35 degrees, in combination with airfoil thickness to chord ratios
equal to or below about 3% as an average along the semi-span
outboard from the zone of substantial fuselage influence to the
wing tip.
Inventors: |
Tracy; Richard R. (Washoe
Valley, NV), Chase; James D. (Reno, NV) |
Assignee: |
Aerion Corporation (Reno,
NV)
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Family
ID: |
46672894 |
Appl.
No.: |
12/932,091 |
Filed: |
February 16, 2011 |
Prior Publication Data
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Document
Identifier |
Publication Date |
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US 20120043430 A1 |
Feb 23, 2012 |
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Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
Issue Date |
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12589424 |
Oct 26, 2009 |
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Current U.S.
Class: |
244/35R |
Current CPC
Class: |
B64C
3/10 (20130101); B64C 30/00 (20130101); B64C
3/14 (20130101); Y02T 50/10 (20130101) |
Current International
Class: |
B64C
3/14 (20060101) |
Field of
Search: |
;244/35R,35A,36,198,204,45R |
References Cited
[Referenced By]
U.S. Patent Documents
Foreign Patent Documents
Other References
Peter Sturdza, Valerie M. Manning, Ilan M. Kroo, Richard R. Tracy,
"Boundary Layer Calculations Preliminary Design of Wings in
Supersonic Flow", American Institute of Aeronautics and
Astronautics, pp. 1-11, USA. cited by other.
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Primary Examiner: Dinh; Tien
Assistant Examiner: Dixon; Keith L
Attorney, Agent or Firm: Haefliger; William W.
Parent Case Text
This is a continuation in part of our prior application, LAMINAR
FLOW WING OPTIMIZED FOR SUPERSONIC CRUISE AIRCRAFT Ser. No.
12/589,424, filed Oct. 26, 2009.
Claims
What is claimed is:
1. An aircraft having maximum efficient cruise speed in a range
from about Mach 0.8 to about Mach 1.2, and having a fuselage and
wings with: (a) less than about 25 degrees of leading edge sweep,
in combination with airfoil thickness to chord ratios between about
3% and about 8%, as an average along a wing semi-span outboard from
a zone of substantial fuselage influence, and (b) a wing leading
edge sweep between about 20 degrees and about 35 degrees, in
combination with airfoil thickness to chord ratios equal to or
below about 3% as an average along the semi-span outboard from the
zone of substantial fuselage influence to a wing tip.
2. The aircraft of claim 1 including fuselage lengthwise contours
to minimize or reduce a wave drag of a combined wing and fuselage,
including engine nacelles and empennage and other lifting and
control surfaces.
3. The aircraft of claim 2 having fuselage and wing contours,
including airfoil and sweep, to minimize or reduce total drag of
the combined wing-fuselage, including effects of laminar flow in
reducing skin friction drag.
4. The aircraft of claim 3 having empennage contours to minimize or
reduce total drag of a combined wing, fuselage and empennage.
5. The aircraft of claim 4 having wing and empennage contours to
maximize or increase a total range of the aircraft for a given
maximum total takeoff weight or other related constraint, including
the effects of thickness-to-chord ratio and sweep on a structural
weight of the wings and empennage.
6. The aircraft of claim 4 having wing and empennage contours to
minimize or decrease a total weight of the aircraft for a given
mission or mix of missions, including effects of thickness-to-chord
ratio and sweep on a structural weight of the wings and
empennage.
7. The aircraft of claim 4 having wing and empennage contours to
minimize or decrease a total weight of the aircraft for a given
anticipated flight mission or mix of missions, including effects of
thickness-to-chord ratio and sweep on a structural weight of the
wings and empennage, where such wing and empennage contours are
characterized by either (a) an iterative process combining
configuration design experience, aerodynamic analysis and wind
tunnel testing, or (b) a computer-based multi-disciplinary
optimization method combined with configuration design
constraints.
8. The aircraft of claim 1 wherein a drag rise Mach number lies
between about 0.80 and 0.97.
9. The aircraft of claim 1 wherein a drag rise Mach number lies
between about 0.97 and 1.2.
10. The aircraft of claim 1 wherein a drag rise Mach number lies
between about 0.92 and 0.97, for a wing average t/c of about
3%.
11. The aircraft of claim 1 wherein a drag rise Mach number lies
between about 0.95 and 1.15, for a wing average t/c of about
3%.
12. The aircraft of claim 1 having about 3% average wing t/c,
characterized in that Mdd, corresponding to about 0.002 drag
coefficient increase, occurs at about Mach 0.96.
13. The aircraft of claim 1 having about 3% average wing t/c,
characterized in that wing leading edge sweep is between about 20
degrees and about 35 degrees, and by a supersonic drag rise Mach
number between about 0.99 and 1.05.
14. The aircraft of claim 1 having wing leading edge sweep less
than about 25 degrees, characterized in that wing average t/c is
between about 3% and 8%.
15. The aircraft of claim 1 having wing leading edge sweep between
about 20 degrees and 35 degrees, characterized in that wing average
t/c is less than about 3%.
Description
This invention relates generally to the configuration of transonic
aircraft with wings designed for extensive natural laminar flow
(NLF), and more particularly to optimization of wing thickness,
sweep and fuselage cross section relationship criteria, for such
transonic aircraft.
BACKGROUND OF THE INVENTION
Transonic NLF wing aircraft configurations as described herein are
desirable for efficient transonic cruise, e.g. high subsonic speeds
typically above about Mach 0.80 and up to slightly above Mach 1.
Principal features of the herein described configurations are low
to moderate sweep, sharp or slightly blunted leading edge, and
relatively thin airfoils in terms of the ratio of maximum airfoil
thickness to chord (t/c). The importance of the NLF boundary layer
(BL) in terms of drag reduction can be understood by considering
that for typical transonic cruise flight conditions the laminar
skin friction drag is approximately 10% of the turbulent skin
friction drag associated with a traditional swept wing designs, for
the same amount of surface area. Additionally, the transonic NLF
wing configurations described herein can achieve best cruise
efficiency at higher Mach numbers than possible with the swept
wings typically used on high speed subsonic cruise aircraft.
For extensive NLF, the wing must have low or moderate sweep, and
thus, on a purely aerodynamic basis the low sweep NLF wing should
be relatively thin to limit the volume wave drag at the design
cruise Mach number. On the other hand a thinner wing incurs a
weight penalty, since structural weight varies inversely with wing
thickness, everything else being equal, so that selection of
thickness to chord ratio (t/c) is of substantial importance to
optimizing the performance of such aircraft.
In previous studies, the NLF wing was designed to give best
efficiency at speeds of about Mach 0.95 or higher. This work formed
certain bases for U.S. Pat. No. 7,000,870, "LAMINAR FLOW WING FOR
TRANSONIC CRUISE", incorporated herein by reference. This Mach
number criterion led to the provision of about 3% (0.03) as an
upper limit for the span-wise average t/c ratio of the NLF wing and
leading edge sweep angles of less than about 20.
However, that prior patent specified no variation of t/c with
design cruise Mach number, M. Design studies have been extended to
cover a range of cruise Mach numbers down to about Mach 0.80, near
the maximum efficient cruise Mach number of previous subsonic
aircraft with low wing sweep designed for long range. These studies
showed that low or moderately swept NLF wings having average t/c up
to about 0.08 (8%) would fill a gap in efficient cruise Mach number
between about 0.80 and about Mach 0.95. Such wings can be designed
for extensive NLF by methods described in our prior patents and the
patent application referenced above (Ser. No. 12/589,424) of which
this is a continuation in part.
In addition, certain design combinations of wing sweep and t/c, can
enable efficient cruise Mach numbers up to about 1.05, well beyond
the maximum efficient cruise Mach number of high speed, long range
aircraft other than supersonic designs capable of operating at more
than about Mach 1.2. Such wings were found to require average t/c
ratios of about 0.03 (3%) or less, and for some missions could
benefit from greater leading edge sweep than the previous limit of
about 20 degrees specified in our U.S. Pat. No. 7,000,870. For
example a sweep of about 30 degrees is required for an efficient
cruise Mach number of 0.99 with an average t/c ratio of about 0.03
(3%). Achieving extensive NLF for such wing sweep is more difficult
and some loss in LF coverage extent is inevitable.
We have found the foregoing combinations of thickness and sweep to
be advantageous for efficient flight at transonic speeds and
determined that these combinations have not been used or disclosed
previously. For example there are many subsonic aircraft which are
limited to maximum cruise speeds of less than Mach 0.80, and which
utilize low sweep, but have thicker wings than the herein proposed
t/c upper limit of 8%. On the other hand there are aircraft such as
commercial jet airliners and high speed business jets, which are
designed for efficient cruise speeds above Mach 0.8 but which have
much higher than 25 degrees of wing sweep, or t/c greater than 8%.
Finally there are actual and proposed supersonic aircraft such as
"Concorde", designed for cruise speeds well above Mach 1.2, which
feature t/c below 3%, but use very high leading edge sweep greater
than about 50 degrees.
As previously noted in prior application (Ser. No. 12/589,424)
titled, LAMINAR FLOW WING OPTIMIZED FOR SUPERSONIC CRUISE AIRCRAFT,
a number of considerations may drive the optimal thickness to
higher values, even at the expense of a moderate increase in volume
wave drag for a given design Mach number. For example the favorable
pressure gradient, which stabilizes the laminar boundary layer,
increases with t/c ratio, and as noted, structural weight decreases
with increasing thickness. In addition, the volume wave drag
attributable to the wing can be reduced by contouring the fuselage
in the vicinity of the wing. Finally, the achievement of NLF on
large areas of the wing surface is dependent on (a) achieving
appropriate pressure gradients over the affected surfaces of the
wing and (b) suitable leading edge size and shape. These pressure
gradients depend not only on the local airfoil shapes, but also are
influenced by the fuselage contour or contours adjacent to the
wing.
There is, accordingly, a need for improvements in cruise efficiency
and range of transonic aircraft, and particularly in the
optimization of the airfoil shapes, thickness to chord ratios, wing
sweep and aspect ratio, as well as the fuselage contours affecting
both volume wave drag and NLF extent over the wing surfaces.
Similar considerations can be applied to the design of horizontal
and vertical tail surfaces.
SUMMARY OF THE INVENTION
The present invention extends the use of wing configurations for
aircraft designed for efficient cruise at transonic speed,
described in prior U.S. Pat. No. 7,000,870, "LAMINAR FLOW WING FOR
TRANSONIC CRUISE" and in patent application Ser. No. 12/589,424,
LAMINAR FLOW WING OPTIMIZED FOR SUPERSONIC CRUISE AIRCRAFT, of
which this is a continuation in part, to aircraft designed for
maximum operating Mach number (Mmo) of greater than about 0.80 up
to about Mach 1.2, as follows: (a) wings having less than about 25
degrees leading edge sweep, in combination with increased average
thickness to chord ratios from about 3% up to about 8% as an
average along the wing span outboard of the zone of fuselage
influence, and (b) having more than about 20 degrees leading edge
sweep up to about 35 degrees, in combination with thickness to
chord ratios less than about 3% as an average along the wing span
outboard of the zone of fuselage influence.
The wing thickness to chord ratios may typically vary from outboard
of the zone of substantial fuselage influence to the wing tip, such
that the average of such ratio along such portion of the span is
consistent with the criteria stated in the foregoing. Wing sweep
refers to the leading edge of the basic "trapezoidal wing", or to
the minimum leading edge sweep in other cases such as the "ogive"
wing plan or wings with an inboard strake. The invention includes
all airfoil types such as bi-convex, NACA 6-series and
supercritical, appropriately modified to provide, at design cruise
Mach number and lift coefficient, an optimal combination of (a)
extensive laminar flow over upper, lower, or both surfaces, (b) low
wave drag, and also (c) high lift at low speed.
The fuselage is preferably contoured to reduce or minimize wave
drag of the wing-fuselage combination, including engine nacelles
and empennage. The foregoing principals generally apply to tail
surfaces as well. Such wing and fuselage contours to achieve
optimal mission performance can be accomplished by either (a) an
iterative process combining configuration design experience,
aerodynamic analysis and wind tunnel testing, or (b) a
computer-based multi-disciplinary optimization method combined with
configuration design constraints. The foregoing design processes
also would preferably include propulsion nacelles and tail or other
lifting or control surfaces such as canards. The shaping
optimization also would preferably include the effect of the
distribution of t/c on wing weight and thus on range at a given
overall maximum weight or maximum weight for a given range.
DRAWING DESCRIPTION
FIG. 1 shows such an aircraft in plan view, as well as a cross
section of the wing showing a typical airfoil in FIG. 1a.
FIG. 2 shows a representative transonic aircraft configuration.
FIG. 3 is a graph of the pressures on the upper and lower surfaces
of an airfoil at two Mach numbers.
FIG. 4 is a graph showing a relationship between airfoil t/c and
drag-rise Mach number.
FIG. 5 is a graph showing the drag of a wing-fuselage configuration
with 3% average wing t/c.
FIG. 6, is a graph of showing higher critical Mach number as a
function of wing leading edge sweep.
FIG. 7 is a chart showing the combination of wing sweep and t/c
ratio for various representative aircraft.
DETAIL DESCRIPTION
FIG. 1 shows an aircraft 9 incorporating the invention including a
fuselage 5, a jet engine nacelle 6 including inlet and exhaust ends
6a and 6b, tail 7 and wing 1. An integrated fuselage/nacelle is
illustrated, but the invention applies also to aircraft with
separate engine nacelles mounted on the wing or fuselage.
The wing leading edge sweep angle ^ is defined as the minimum angle
of the outboard trapezoidal wing leading edge 2 relative to a line
projected normally outboard from the aircraft longitudinal axis.
FIG. 1a is a chordwise vertical section A-A through the wing 8, and
is generally representative of the wing t/c ratio, where these
dimensions are shown in section A-A. For the present purposes, the
wing t/c is defined as the average of the t/c values along the wing
span from a location outboard of the zone of appreciable fuselage
influence on wing drag to the wing tip 4.
Location 11 shows a reduction in cross-sectional area of the
fuselage and/or nacelle adjacent to wing 1 to reduce wave drag in
accordance with area rule principals, as well as to reduce viscous
drag by suppressing cross-flow pressure gradients across the wing
surface, which are generally adverse to laminar flow. Location 12
shows a similar modification of fuselage cross-sectional area
adjacent to tail 7.
FIG. 2 shows an alternative preferred aircraft configuration 20
incorporating the invention, featuring a fuselage body 22, engine
nacelle 23, wing 21, aft located horizontal tail surface 24, aft
located vertical tail surface 26, and forwardly located canard
surface 25. Any or all of the wing and stabilizing or control
surfaces 21, 24, 25, 26 may incorporate leading edge sweep and
thickness geometries as described previously. The fuselage 22
and/or nacelle 23 may have reduced cross-section or "waisting"
adjacent to surfaces 21, 24, 25, or 26 to reduce drag as described
above.
FIG. 3 illustrates the pressures on the upper and lower surfaces of
an airfoil at two Mach numbers. It shows the effect of higher
subsonic Mach number in creating pressure conditions favorable to
laminar flow, namely the negative pressure gradients (pressure
decreasing with distance aft) over the majority of both upper and
lower surfaces at Mach 0.95. For comparison the positive pressure
gradients at Mach 0.8, especially on the upper surface, are adverse
to extensive runs of laminar flow.
FIG. 4 is a graph showing a representative relationship between
airfoil t/c and the drag divergence Mach number (Mdd), for which
the increase in drag coefficient is 0.002. The data is calculated
using high order aerodynamic codes able to correctly represent
conditions near Mach 1. The airfoils for this example are
sharp-edged bi-convex sections at zero lift, and thus represent the
lowest drag at a given Mach number and t/c. Airfoils with blunt
leading edges, camber and angle of attack corresponding to a
representative cruise lift coefficient will have a somewhat lower
drag divergence Mach number, but the graph is illustrative of the
relationship of Mdd to t/c.
FIG. 5 illustrates the relatively low drag-rise at high subsonic
Mach number of a wing-fuselage configuration with 3% average wing
t/c, both at zero lift and at a lift coefficient of 0.3, typical of
transonic cruise conditions. In both cases the drag divergence Mach
number occurs at about Mach 0.96, much higher than achieved to
applicants' knowledge with any current conventional swept wing
subsonic aircraft.
FIG. 6 shows the approximate increase in critical Mach number (a
widely used criterion for incipient drag rise with increasing Mach
number) as a function of quarter-chord sweep for wings typical of
those used in industry. This curve can be considered a first-order
estimate of the role of leading edge sweep in increasing the low
drag Mach number for a given t/c and airfoil. For example, the
figure shows that a sweep angle of 30 degrees should increase Mdd
by about 0.06. But from FIG. 3, an unswept wing with 3% t/c would
have an Mdd of about 0.93, thus the 30 degree swept wing with 3%
t/c would be expected to achieve efficient low drag flight at about
Mach 0.99. This result would vary depending on details of the
airfoil, wing span and integration with a fuselage, but is a much
higher drag rise Mach number than can achieved to applicants'
knowledge with any conventional 30 degree swept wing.
FIG. 7 is a chart showing the combination of wing sweep and wing
t/c ratio for all representative aircraft with maximum operating
cruise speeds, Mmo, of more than Mach 0.80. Only selected points
are identified as to the specific aircraft models to reduce clutter
and emphasize known aircraft.
The chart shows that all aircraft with wing sweep angles below
about 25 degrees, have thicker wings than the proposed 8% t/c upper
limit. There are numerous subsonic aircraft such as jet transports
and high end business jets designed for Mmo greater than Mach 0.80,
but none having wing sweep less than 20 degrees and t/c below 8%.
Also there are a few supersonic aircraft, mainly fighters, designed
for Mmo up to Mach 2 or more, employing average t/c greater than
about 0.03, but with wing sweep well above 20 degrees. Finally
there are long-range supersonic cruise aircraft (or published
designs) with average t/c less than 0.03 but all having leading
edge sweep greater than about 50 degrees.
Thus, based on our comprehensive research, as FIG. 7 indicates,
there are no aircraft that embody the proposed configuration
combinations, namely (a) leading edge sweep less than about 25
degrees combined with average t/c between about 3% and 8%, and (b)
leading edge sweep between about 20 and 35 degrees combined with
average t/c less than about 0.03 (3%).
* * * * *